A hybrid vane ( 50 ) for a gas turbine engine having a ceramic matrix composite (CMC) airfoil member ( 52 ) bonded to a substantially solid core member ( 54 ). The airfoil member and core member are cooled by a cooling fluid ( 58 ) passing through cooling passages ( 56 ) formed in the core member. The airfoil member is cooled by conductive heat transfer through the bond (( 70 ) between the core member and the airfoil member and by convective heat transfer at the surface directly exposed to the cooling fluid. A layer of insulation ( 72 ) bonded to the external surface of the airfoil member provides both the desired outer aerodynamic contour and reduces the amount of cooling fluid required to maintain the structural integrity of the airfoil member. Each member of the hybrid vane is formulated to have a coefficient of thermal expansion and elastic modulus that will minimize thermal stress during fabrication and during turbine engine operation.

Ceramic matrix composite gas turbine vane

A variable geometry turbocharger includes vanes disposed therein having an inner airfoil surface oriented adjacent a turbine wheel, and an outer airfoil surface oriented opposite and parallel to the inner airfoil surface. The vane includes first and second axial surfaces that are each positioned perpendicular to and interposed between the inner and outer airfoil surfaces. A vane leading edge is positioned along a first inner and outer airfoil surface junction, and a trailing edge is positioned along a second inner and outer airfoil surface junction. One or more of the first and second axial surfaces have a composite construction comprising a solid section that extends a distance from the leading edge towards the trailing edge, and a cored-out section that extends a distance from the trailing edge towards the leading edge.

Variable geometry turbocharger

A method of manufacturing a composite structure uses a layer of an insulating material ( 22 ) as a mold for forming a substrate of a ceramic matrix composite (CMC) material ( 24 ). The insulating material may be formed in the shape of a cylinder ( 10 ) with the CMC material wound on an outer surface ( 14 ) of the cylinder to form a gas turbine combustor liner ( 20 ). Alternatively, the insulating material may be formed in the shape of an airfoil section ( 32 ) with the CMC material formed on an inside surface ( 36 ) of the insulating material. The airfoil section may be formed of a plurality of halves ( 42, 44 ) to facilitate the lay-up of the CMC material onto an easily accessible surface, with the halves then joined together to form the complete composite airfoil. In another embodiment, a box structure ( 102 ) defining a hot gas flow passage ( 98 ) is manufactured by forming insulating material in the shape of opposed airfoil halves ( 104 ) joined at respective opposed ends by platform members ( 109 ). A layer of CMC material ( 107 ) is then formed on an outside surface of the insulating material. A number of such composite material box structures are then joined together to form a vane ring ( 100 ) for a gas turbine engine.

Composite structure formed by cmc-on-insulation process

Aspects of the invention relate to a modular turbine vane assembly. The vane assembly includes an airfoil portion, an outer shroud and an inner shroud. The airfoil portion can be made of at least two segments. Preferably, the components are connected together so as to permit assembly and disassembly of the vane. Thus, in the event of damage to the vane, repair involves the replacement of only the damaged subcomponents as opposed to the entire vane. The modular design facilitates the use of various materials in the vane, including materials that are dissimilar. Thus, suitable materials can be selected to optimize component life, cooling air usage, aerodynamic performance, and cost. Because the vane is an assemblage of smaller sub-components as opposed to one unitary structure, the individual components of the vane can be more easily manufactured and more intricate features can be included.

Multi-piece turbine vane assembly

A vane has an airfoil shell and a spar within the shell. The vane has an outboard shroud at an outboard end of the shell and an inboard platform at an inboard end of the shell. The shell includes a region having a depth-wise coefficient of thermal expansion and a second coefficient of thermal expansion transverse thereto, the depth-wise coefficient of thermal expansion being greater than the second coefficient of thermal expansion.

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Ceramic matrix composite turbine engine vane

A vane assembly for a turbine assembly includes an inner endcap, an outer endcap, and a body. The body includes a metallic core assembly, a ceramic shell assembly and a support assembly. The metallic core assembly is coupled to the inner and outer endcaps and bears most of the mechanical loads, including aerodynamic loads. The ceramic shell bears substantially all of the thermal stress placed on the vane assembly. The support assembly is disposed between the metallic core assembly and said ceramic shell assembly and is coupled to the metallic core assembly.

Ceramic composite vane with metallic substructure

Embodiments of the invention relate to a robust turbine vane made of stacked airfoil-shaped CMC laminates. Each laminate has an in-plane direction and a through thickness direction substantially normal to the in-plane direction. The laminates have anisotropic strength characteristics in which the in-plane tensile strength is substantially greater than the through thickness tensile strength. Thus, the laminates can provide strength in the direction of high thermal gradients and, thus, withstand the associated high thermal stresses. The laminates are relatively weak in through thickness (interlaminar) tension, but, in operation, relatively low through thickness tensile stresses can be expected. The laminates can be strong in through thickness compression; accordingly, the laminate stack can be held in through thickness compression by one or more fasteners. The CMC material can permit the inclusion of additional features such as cooling passages, ribs, spars, and thermal coatings, without compromising the strength characteristics of the material.

Stacked laminate CMC turbine vane

An article comprising a ceramic material having a ceramic matrix composite backing adapted for use in a gas turbine engine is provided. The article comprises a structural ceramic material having a hot side facing toward a high temperature environment and a cold side facing away from the high temperature environment; and a ceramic matrix composite composition having a strength greater than the strength of the ceramic material attached to the back of the cold side of the ceramic material, whereby crack initiation and propagation are inhibited by the ceramic matrix composition to a greater degree than by the ceramic material.

Ceramic material having ceramic matrix composite backing and method of manufacturing

Embodiments of the invention relate to a turbine vane having a fail safe cooling system. According to embodiments of the invention, the vane can have multiple concentric layers of radial cooling holes extending about the vane; each layer being fluidly connected to the adjacent layer or layers. Such fluid communication can occur through one or more plenums in the vane or in the shrouds bounding the radial ends of the vane. Coolant can initially be supplied to the innermost layer of cooling holes. From there, the coolant can sequentially progress through successive outer layers. Between two adjacent layers, the coolant can flow in opposite directions. Not only does such a system provide needed cooling to the vane, but the multilayer redundant cooling system can avoid or delay catastrophic failures that can occur if the vane surface is damaged, such as by impact.

Daniel G. Thompson

Papers

Ceramic matrix composite gas turbine vane

Ceramic composite vane with metallic substructure

Stacked laminate CMC turbine vane

Ceramic material having ceramic matrix composite backing and method of manufacturing

Fail safe cooling system for turbine vanes